This invention relates to gas turbine engines having bypass airflow with a variable exhaust nozzle suitable for aircraft use and, more particularly, to the control of engine operation based on ratios of pressure at different locations within the engine.
One class of turbine engine for aircraft use which is of particular interest herein is known as the "variable cycle" engine. The variable cycle is a family of hybrid gas turbine engines which can operate with the high specific thrust characteristics of a low bypass-ratio turbofan or turbojet engine at supersonic speeds, yet also exhibit the lower specific thrust, low noise and low fuel consumption characteristics of a high bypass-ratio turbofan engine. The need for such variable cycle engines has risen because of the need for an efficient multimission aircraft.
There has been a need to establish the best control mode for a variable cycle engine. The control mode should obtain the most accurate control or regulation of those factors which affect engine operability, life and performance. These factors are net thrust, high-pressure turbine inlet temperature, fan stall margin, booster-stage stall margin, maximum rotor speeds, and cooling-flow pressure ratio. This regulation should be maintained with minimal effects from errors of control sensors, variations in engine quality, and engine deterioration from wear or service. The foregoing characteristics are most critical at the maximum dry power and at augmented thrust levels above the maximum dry power.
Experience has shown that the principle factor in achieving the foregoing control regulation objective is the selection of a set of measured engine variables that are the most effective and mutually compatible. These control parameters may include a variable engine geometry such as nozzle throat cross section, or stator position, by way of example, as well as a pressure ratio, a rotor speed, or a rotor speed "corrected" by a measured temperature. An engine may have several control modes (sets of control variables) with each mode being used for a different power range or flight operating region, such as altitude and Mach number ranges.
A major advancement in engine-cycle design, or the addition of new engine elements of variable geometry, usually requires a reconsideration of the control modes relative to those used in earlier or predecessor engines. This presents a problem in that the control modes used for one family of engines need to be changed, at least in part, to meet the special needs of the later improved engine. The previously defined modes are not necessarily fully satisfactory in the regulating of net thrust and other parameters, mentioned above, for the later engine. Other control modes and parameters which have been practiced in the construction and use of aircraft turbine engines have not met all of the foregoing goals. Examples of such construction and practice which may not be satisfactory for meeting the foregoing goals are the following: (A) Turbine exhaust gas temperature or computed turbine inlet temperature controls may cause excessive thrust loss when considering the effects of component deterioration. (B) Optical pyrometers for turbine metal temperatures may present problems in terms of cost and reliability. (C) Engine pressure ratio (turbine discharge pressure divided by fan-inlet pressure) has the disadvantage of requiring the measurement of engine-inlet pressure in a highly distorted pressure field subject to serious acquisition errors, as may be caused by a poor sample of inlet pressure. Also, icing of pressure probes or orifices, as well as of pressure lines should be prevented. Presently available technology has not met adequately the foregoing goals.